Oil system bearing compartment architecture for gas turbine engine

ABSTRACT

A gas turbine engine with a geared architecture includes a multiple of bearing compartments and at least one carbon seal that seals at least one side of each of the multiple of bearing compartments.

This application claims priority to U.S. Provisional Application No.61/719,162 filed Oct. 26, 2012, which is hereby incorporated byreference.

BACKGROUND

The present disclosure relates to a gas turbine engine and, moreparticularly, to an oil system bearing compartment arrangement therefor.

Aircraft gas turbine engines include an oil system to supply oil tovarious components such as bearings that are typically contained withinbearing compartments. All gas turbine engines ingest some air from thegas path in regular operation due to the pressure differential betweenthe gas path and bearing compartments. Under certain conditions, the oilmay be churned at a high velocity and thereby become aerated. If the oilis not quieted and deaerated, the oil may not be effectively scavenged.In a geared engine architecture, higher air pressures within bearingcompartments that contain a gear system can reduce the efficiency of thegear system. Entrained air in the oil is removed through a deaerator toassure quality oil, however, removal of air may also result in someundesirable oil consumption. The oil must eventually be replaced duringmaintenance operations.

SUMMARY

A gas turbine engine with a geared architecture according to onedisclosed non-limiting embodiment of the present disclosure includes amultiple of bearing compartments and at least one carbon seal whichseals at least one side of each of the multiple of bearing compartments.

In a further embodiment of the foregoing embodiment, the multiple ofbearing compartments include a front bearing compartment.

In a further embodiment of any of the foregoing embodiments, themultiple of bearing compartments include a mid bearing compartment.

In a further embodiment of any of the foregoing embodiments, themultiple of bearing compartments include a mid-turbine bearingcompartment.

In a further embodiment of any of the foregoing embodiments, themultiple of bearing compartments include a rear bearing compartment.

In a further embodiment of any of the foregoing embodiments, each of themultiple of bearing compartments interface with an engine shaft.

In a further embodiment of any of the foregoing embodiments, the gearedarchitecture is 98% efficient.

In a further embodiment of any of the foregoing embodiments, themultiple of bearing compartments include a front bearing compartment, amid bearing compartment axially aft of the front bearing compartment, amid-turbine bearing compartment axially aft of the mid bearingcompartment, and a rear bearing compartment axially aft of themid-turbine bearing compartment.

A gas turbine engine according to another disclosed non-limitingembodiment of the present disclosure includes a front bearingcompartment bounded by a first and second carbon seal, a mid bearingcompartment bounded by a first and second carbon seal, the mid bearingcompartment axially aft of the front bearing compartment, a mid-turbinebaring compartment bounded by a first and second carbon seal, themid-turbine bearing compartment axially aft of the mid bearingcompartment, and a rear bearing compartment bounded by a first andsecond carbon seal, the rear bearing compartment axially aft of themid-turbine bearing compartment.

In a further embodiment of the foregoing embodiment, the front bearingcompartment contains #1 and a #1.5 tapered bearing and a gearedarchitecture which drives a fan at a lower speed than a low spool.

In a further embodiment of any of the foregoing embodiments, themid-bearing compartment contains a #2 bearing that rotationally supportsa forward end section of a low spool and a #3 bearing that rotationallysupports a forward end section of a high spool.

In a further embodiment of any of the foregoing embodiments, themid-turbine bearing compartment contains a #4 bearing that rotationallysupports an aft end section of a high spool.

In a further embodiment of any of the foregoing embodiments, the rearbearing compartment contains a #5 bearing and a #6 bearing thatrotationally supports an aft end section of a low spool.

In a further embodiment of any of the foregoing embodiments, furthercomprises an oil system in fluid communication with each of the frontbearing compartment, the mid bearing compartment, the mid-turbinebearing compartment and the rear bearing compartment. In the alternativeor additionally thereto, the foregoing embodiment includes the oilsystem includes a deoiler. In the alternative or additionally thereto,the foregoing embodiment includes the oil system includes a deaerator.

A method of reducing oil outflow from an oil system in communicationwith a geared architecture of a gas turbine engine having a multiple ofbearing cavities, according to another disclosed non-limiting embodimentof the present disclosure includes bounding a multiple of bearingcavities with a multiple of carbon seals.

In a further embodiment of the foregoing embodiment, further comprisesbounding each of the multiple of bearing cavities with a first andsecond carbon seal.

In a further embodiment of any of the foregoing embodiments, furthercomprises associating each of the multiple of bearing cavities with aspool of the gas turbine engine.

In a further embodiment of any of the foregoing embodiments, one of thebearing cavities contains a geared architecture which drives a fan at arotational speed less than a low spool.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-sectional view of a geared architecture gasturbine engine;

FIG. 2 is a schematic view of an oil system for the geared architecturegas turbine engine;

FIG. 3 is a graphical representation of seal air leakage in standardcubic feet of air per minute (SCFM) on the ordinate and differentialpressure (deltaP) across the seal on the abscissa; and

FIG. 4 is an expanded cross-sectional view of a carbon seal according toone disclosed non-limiting embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative enginesarchitectures such as a low-bypass turbofan may include an augmentorsection (not shown) among other systems or features. Althoughschematically illustrated as a turbofan in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with turbofans as the teachings may be applied toother types of turbine engines to include but not limited to athree-spool (plus fan) engine wherein an intermediate spool includes anintermediate pressure compressor (IPC) between a low pressure compressorand a high pressure compressor with an intermediate pressure turbine(IPT) between a high pressure turbine and a low pressure turbine as wellas other engine architectures such as turbojets, turboshafts, openrotors and industrial gas turbines.

The fan section 22 drives air along a bypass flowpath and a coreflowpath while the compressor section 24 drives air along the coreflowpath for compression and communication into the combustor section 26then expansion through the turbine section 28. The engine 20 generallyincludes a low spool 30 and a high spool 32 mounted for rotation aboutan engine central longitudinal axis A relative to an engine caseassembly 36 via several bearing compartments 38-1, 38-2, 38-3, 38-4. Thelow spool 30 generally includes an inner shaft 40 that interconnects afan 42, a low pressure compressor 44 (“LPC”) and a low pressure turbine46 (“LPT”). The inner shaft 40 drives the fan 42 through a gearedarchitecture 48 to drive the fan 42 at a lower speed than the low spool30.

The high spool 32 includes an outer shaft 50 that interconnects a highpressure compressor 52 (“HPC”) and high pressure turbine 54 (“HPT”). Acombustor 56 is arranged between the HPC 52 and the HPT 54. The innershaft 40 and the outer shaft 50 are concentric and rotate about theengine central longitudinal axis A that is collinear with theirlongitudinal axes.

Core airflow is compressed by the LPC 44 then the HPC 52, mixed with thefuel and burned in the combustor section 56, then expanded over the HPT54 and the LPT 46. The HPT 54 and the LPT 46 drive the respective lowspool 30 and high spool 32 in response to the expansion.

In one non-limiting example, the gas turbine engine 20 is a high-bypassgeared architecture engine in which the bypass ratio is greater thanabout six (6:1). The geared architecture 48 can include an epicyclicgear train, such as a planetary gear system, star gear system or othergear system. The example epicyclic gear train has a gear reduction ratioof greater than about 2.3, and in another example is greater than about2.5 with a gear system efficiency greater than approximately 98%. Thegeared turbofan enables operation of the low spool 30 at higher speedswhich can increase the operational efficiency of the low pressurecompressor 44 and low pressure turbine 46 and render increased pressurein a fewer number of stages.

A pressure ratio associated with the LPT 46 is pressure measured priorto the inlet of the LPT 46 as related to the pressure at the outlet ofthe LPT 46 prior to an exhaust nozzle of the gas turbine engine 20. Inone non-limiting embodiment, the bypass ratio of the gas turbine engine20 is greater than about ten (10:1), the fan diameter is significantlylarger than that of the low pressure compressor 44, and the low pressureturbine 46 has a pressure ratio that is greater than about five (5:1).It should be understood, however, that the above parameters are onlyexemplary of one embodiment of a geared architecture engine and that thepresent disclosure is applicable to other gas turbine engines includingdirect drive turbofans.

In one non-limiting embodiment, a significant amount of thrust isprovided by the bypass flow due to the high bypass ratio. The fansection 22 of the gas turbine engine 20 is designed for a particularflight condition—typically cruise at about 0.8 Mach and about 35,000feet. This flight condition, with the gas turbine engine 20 at its bestfuel consumption, is also known as bucket cruise Thrust Specific FuelConsumption (TSFC). TSFC is an industry standard parameter of fuelconsumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The low FanPressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed isthe actual fan tip speed divided by an industry standard temperaturecorrection of (“T”/518.7)^(0.5). in which “T” represents the ambienttemperature in degrees Rankine. The Low Corrected Fan Tip Speedaccording to one non-limiting embodiment of the example gas turbineengine 20 is less than about 1150 fps (351 m/s).

The bearing compartments 38-1, 38-2, 38-3, 38-4 in the disclosednon-limiting embodiment are defined herein as a front bearingcompartment 38-1, a mid-bearing compartment 38-2 axially aft of thefront bearing compartment 38-1, a mid-turbine bearing compartment 38-3axially aft of the mid-bearing compartment 38-2 and a rear bearingcompartment 38-4 axially aft of the mid-turbine bearing compartment38-3.

Each of the bearing compartments 38-1, 38-2, 38-3, 38-4 include one ormore bearings 60 (illustrated schematically) and one or more—typicallytwo (2)—carbon seals 62 (illustrated schematically). In anotherdisclosed non-limiting embodiment, only one carbon seal may be locatedon one side of the bearing compartment 38-1, 38-2, 38-3, 38-4 subject togreatest air leakage—the lower air leakage side could use a labyrinthtype seal.

Various types of carbon seals 62 may be used herewith and the carbonseals 62 contemplated herein include, but are not limited to facecontact seals. The bearings 60 and carbon seals 62 respectively supportand interface with the shafts 40, 50 of the respective low spool 30 andhigh spool 32. The carbon seals 62 operate to seal a “wet” zone from a“dry” zone. In other words, regions or volumes that contain oil may bereferred to as a “wet” zone and an oil-free region may be referred to asa “dry” zone. So, for example, the interior of each bearing compartment38-1, 38-2, 38-3, 38-4 may be referred to as a wet zone that ultimatelycommunicates with an oil sump (not shown) of an oil system (FIG. 2)while the region external thereto may be referred to as a dry zone. Thatis, the bearings 60 support the low spool 30 and the high spool 32 andthe carbon seals 62 separate the “wet” zone from the “dry” zone todefine the boundaries of each bearing compartment 38-1, 38-2, 38-3,38-4.

In the disclosed, non-limiting embodiment, the front bearing compartment38-1 contains a #1 tapered bearing 60, a #1.5 tapered bearing 60 and thegeared architecture 48. The #1 tapered bearing 60 and the #1.5 taperedbearing 60 rotationally support the fan 42. In the disclosed,non-limiting embodiment, the mid-bearing compartment 38-2 contains a #2bearing 60 and a #3 bearing 60. The #2 bearing 60 rotationally supportsa forward end section of the low spool 30 and the #3 bearing 60rotationally supports a forward end section of the high spool 32. Themid-turbine bearing compartment 38-3 contains a #4 bearing 60 thatrotationally supports the aft end section of the high spool 32. In thedisclosed, non-limiting embodiment, the rear bearing compartment 38-4contains a #5 bearing 60 and a #6 bearing 60 that rotationally supportan aft end section of the low spool 30. Although particular bearingcompartments and bearing arrangements are illustrated in the disclosednon-limiting embodiment, other bearing compartments and bearingarrangements in other engine architectures such as three-spoolarchitectures will also benefit herefrom.

With reference to FIG. 2, an oil system 70 provides oil under pressureto lubricate and cool moving components of the engine 20, such as, forexample but not limited to, the geared architecture 48 and the bearings60. The oil system 70 generally includes a pump 72, a main filter 74, apressure relief valve 76, an air-oil cooler 78, a fuel-oil cooler 80, anoil pressure trim orifice 82, an oil strainer 84, a deoiler 86, adeareator 88 and a tank 90. The oil system 70 is but a simplifiedschematic illustration and that other additional or alternativecomponents and subsystems may be included in the oil system 70.

In operation, the pump 72 communicates oil to the main filter 74 thenfiltered oil proceeds to the air-oil cooler 78 and/or the fuel-oilcooler 80 to be cooled. Should the main filter 74 become clogged, thepressure relief valve 76 indicates the clog. The oil pressure trimorifice 82 then regulates the oil flow and returns any oil overage tothe oil tank 90. The oil strainer 84 strains any debris before the oilis communicated to the bearing compartment 38-1, 38-2, 38-3, 38-4 tolubricate, for example, the bearings 60.

Oil is then scavenged from the bearing compartment 38-1, 38-2, 38-3,38-4 by the pump 72 through the deoiler 86 for return to the tank 90.Air is separated from the scavenge oil from the pump 72 by the deaerator88 at the tank 90. Vent air from the bearing compartment 38-1, 38-2,38-3, 38-4 and air from the deaerator 88 is communicated directly to thedeoiler 86 where the air is removed from the oil and rejected overboard.The oil cycle is then repeated.

Geared architecture engines are architected to minimize oil flow, heatrejection and air entrained in the oil as well as oil outflow from theoil system in regular operation. In particular, gear systems operatemore efficiently with less air flow into the bearing compartment becausethe additional air creates windage and churning within the rotatinggears and bearings. This additional loss in gear system efficiencyreduces the overall efficiency of the engine, increasing the amount offuel needed to complete a specified mission. All gas turbine engines,however, ingest some air from the gas path due to the pressuredifferential between the gas path and the bearing compartments. Gearedarchitecture engines typically utilize an oil system that has an oilflow circulation that may be approximately twice that of a traditionaldirect drive turbofan, e.g., 45 gallons per minute versus 25 gallons perminute from a 35 quart oil tank vs. a 28 quart system. As such, oilingestion may be amplified. Minimization of entrained air in the oilwill thereby facilitate reduced consumption of oil, the minimization ofmaintenance requirements and longer mission potential.

Complete utilization of carbons seals 62 significantly reduces the airleakage (FIG. 3) into each bearing compartment 38-1, 38-2, 38-3, 38-4 ascompared to labyrinth seals that are typically utilized to seal bearingcompartments of conventional direct drive turbofan engines. It should beunderstood that all carbon seals form a seal with a smooth carbonsurface rubbing against a smooth metal surface and that various carbonseal types and structures may be utilized herewith.

With reference to FIG. 4, an illustrated embodiment of the carbon seal62 includes a seal race 100, a fixed segment 102, a carbon element 104,a preload 106 and a secondary seal 108. The seal race 100 may be mountedto the shaft 40, 50 for rotation therewith. The seal race 100 may alsoinclude cooling passages 110 (one shown) to receive and direct coolingoil toward the carbon element 104. The fixed segment 102 is mounted tothe case assembly 36 to support and preload the carbon element 104 intocontact with the seal race 100 by the preload 106 such as springcompression, centrifugal force, and air pressure. The secondary seal 108may include, for example, a piston ring, O-ring type, C-ring type, orother interface. The piston ring may be manufactured of a metal alloywhile the O-ring type and the C-ring type may be elastomeric and mayinclude non-metallic materials to provide an effective seal.

The reduction of air leakage due to utilization of carbon seals 62 inall the bearing compartments 38-1, 38-2, 38-3, 38-4 reduces airentrainment in the oil such that oil tank dwell time is concomitantlyreduced to facilitate usage of a relatively small deoiler 86, deareator88 and tank 90. Reduced airflow in bearing compartments that house thegear system such as the front bearing compartment 38-1 improves gearsystem efficiency. Even though the geared architecture turbofan enginemay have oil flow requirements greater than that of a conventionaldirect drive turbofan, complete usage of carbon seals 62 facilitate asignificant weight, cost and complexity savings. Furthermore, a quart ofoil weighs approximately 2 lbs. (907 g), so a size reduction of the oilsystem components facilitates further weight savings.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Theforegoing description is exemplary rather than defined by thelimitations within Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. A gas turbine engine with a geared architecture,comprising: a multiple of bearing compartments; and at least one carbonseal which seals at least one side of each of said multiple of bearingcompartments.
 2. The gas turbine engine as recited in claim 1, whereinsaid multiple of bearing compartments include a front bearingcompartment.
 3. The gas turbine engine as recited in claim 1, whereinsaid multiple of bearing compartments include a mid bearing compartment.4. The gas turbine engine as recited in claim 1, wherein said multipleof bearing compartments include a mid-turbine bearing compartment. 5.The gas turbine engine as recited in claim 1, wherein said multiple ofbearing compartments include a rear bearing compartment.
 6. The gasturbine engine as recited in claim 1, wherein each of said multiple ofbearing compartments interface with an engine shaft.
 7. The gas turbineengine as recited in claim 1, wherein said geared architecture is 98%efficient.
 8. The gas turbine engine as recited in claim 1, wherein saidmultiple of bearing compartments include: a front bearing compartment; amid bearing compartment axially aft of said front bearing compartment; amid-turbine bearing compartment axially aft of said mid bearingcompartment; and a rear bearing compartment axially aft of saidmid-turbine bearing compartment.
 9. A gas turbine engine, comprising: afront bearing compartment bounded by a first and second carbon seal; amid bearing compartment bounded by a first and second carbon seal, saidmid bearing compartment axially aft of said front bearing compartment; amid-turbine bearing compartment bounded by a first and second carbonseal, said mid-turbine bearing compartment axially aft of said midbearing compartment; and a rear bearing compartment bounded by a firstand second carbon seal, said rear bearing compartment axially aft ofsaid mid-turbine bearing compartment.
 10. The gas turbine engine asrecited in claim 9, wherein said front bearing compartment contains #1and a #1.5 tapered bearing and a geared architecture which drives a fanat a lower speed than a low spool.
 11. The gas turbine engine as recitedin claim 9, wherein said mid-bearing compartment contains a #2 bearingthat rotationally supports a forward end section of a low spool and a #3bearing that rotationally supports a forward end section of a highspool.
 12. The gas turbine engine as recited in claim 9, wherein saidmid-turbine bearing compartment contains a #4 bearing that rotationallysupports an aft end section of a high spool.
 13. The gas turbine engineas recited in claim 9, wherein said rear bearing compartment contains a#5 bearing and a #6 bearing that rotationally supports an aft endsection of a low spool.
 14. The gas turbine engine as recited in claim9, further comprising an oil system in fluid communication with each ofsaid front bearing compartment, said mid bearing compartment, saidmid-turbine bearing compartment and said rear bearing compartment. 15.The gas turbine engine as recited in claim 14, wherein said oil systemincludes a deoiler.
 16. The gas turbine engine as recited in claim 14,wherein said oil system includes a deaerator.
 17. A method of reducingoil outflow from an oil system in communication with a gearedarchitecture of a gas turbine engine having a multiple of bearingcavities, comprising: bounding a multiple of bearing cavities with amultiple of carbon seals.
 18. The method as recited in claim 14, furthercomprising: bounding each of the multiple of bearing cavities with afirst and second carbon seal.
 19. The method as recited in claim 14,further comprising: associating each of the multiple of bearing cavitieswith a spool of the gas turbine engine.
 20. The method as recited inclaim 14, wherein one of said bearing cavities contains a gearedarchitecture which drives a fan at a rotational speed less than a lowspool.